The assignee of the instant application has proposed so-called auxiliary power units (APU's) or emergency power units (EPU's) that are driven by air breathing turbines and which are intended to be utilized in aircraft such as so-called "fly by wire" aircraft. One requirement of many such systems is that the same be able to be brought up to operational speed in a very short period of time, frequently as little as 2-3 seconds, over a wide range of altitudes. This requirement is particularly difficult to achieve at high altitudes because of insufficient naturally occurring oxidant at those altitudes. Thus, the assignee has proposed what have been referred to as stored energy systems for use in connection with such APU's or EPU's, and which typically include an auxiliary combustor, that is, a combustor in addition to the usual combustor or combustors employed in the gas turbine itself. The auxiliary combustor is provided with oxidant such as oxygen, oxygen-enriched air, or air from a storage vessel, typically a pressure bottle.
At high altitudes, relatively cold conditions are present and it is not unusual for the mass of the system including the oxidant, fuel to be utilized as well as the metal components of the system to be at temperatures on the order of -40 degrees Fahrenheit. When the oxidant is flowed to the combustor, the Joule-Thompson effect instantaneously lowers the oxidant temperature to a very low level. Temperatures as low as -110 degrees F. have been observed.
Typical fuels have a freezing point of about -50 degrees Fahrenheit with a consequence that the injection of well atomized fuel into the cold oxidant can result in freezing of the fuel before the fuel ignites.
Moreover, as the oxidant in the storage bottles becomes depleted during the operation of the system, an increasing drop in oxidant temperature occurs and has resulted in the oxidant being delivered at a temperature as low as -160 degrees F. In such a situation, even if ignition was initially obtained, the fuel might later freeze during operation and prevent proper operation of the system.
The problem becomes increasingly acute as combustor size is reduced and those familiar with the requirements of airborne systems will readily appreciate that a constant effort is made to minimize size of aircraft components in terms of minimizing both weight and volume.
In the case of an auxiliary combustor for systems of the type of concern, conventional wisdom has dictated that the ignitor or ignitors be located at the radially outer extremity of the combustor chamber. Such a location, it is felt, assures good ignition because of the longer distance the fuel and oxidant must travel from the typical injector location at the chamber inlet before encountering the igniter at the lowest possible velocity. With a small volume combustor, as volume is decreased, the flame recirculation zone becomes progressively smaller. This zone is required to assure continued ignition of incoming fuel and if the same becomes too small, the initial kernel of flame from the igniter may be swept out of the combustor itself without contacting and igniting the recirculating flow. Conventionally locating the igniter accentuates this possibility.
The present invention is directed to overcoming one or more of the above problems.